Current state-of-the-art propulsion systems include multiple-pulse motors that provide a measure of motor burn controllability in the form of liquid propellant thrusters. However, excess fuel flow is required for wall cooling in these liquid rocket engines, resulting in lower propulsive efficiency. Other known propulsion systems include conventional hybrid rocket propulsion systems in which solid fuel burns with a liquid or gaseous oxidizer flowing through the center of a hollow solid propellant grain creating a diffusion flame and providing thrust.
These prior systems have numerous shortcomings that make them undesirable or unsuitable for use. Rockets usually have severe weight limits for components related to propellant storage, thermal management systems, and the motor structure. In addition, rocket bodies, and particularly those classified as miniature devices, have a highly limited volume for total system packaging. Despite these limitations, rockets are primarily used in military and scientific exo-atmospheric applications that still require high-performance. Furthermore, these systems may be prone to undesirable thermal soak, resulting in unwanted initiation of propellant ignition in a multiple-pulse operation.
Thus, what is needed is a rocket propulsion system that meets these weight and volume restrictions while still providing high performance, so that selection of suitable propellants have high specific impulse and the rocket has a high mass fraction of propellant compared to the whole rocket vehicle weight.
What is also needed is a rocket propulsion system that overcomes thermal soak problems that can result in unwanted propellant ignition.